1. Field of the Invention
The present invention relates to a gas turbine engine for an aircraft and such and more particularly relates to a shroud segment of a turbine shroud for the gas turbine engine.
2. Description of the Related Art
A turbine of a gas turbine engine for an aircraft and such is provided with plural stages of turbine shrouds for suppressing influence of hot combustion gas on a turbine case. The turbine shrouds are directly exposed to the hot gas and thereby a severe thermal stress might be applied thereto. To avoid an excessive thermal stress, in common, each turbine shroud is segmented. Plural shroud segments are built up to form each turbine shroud.
Each shroud segment is provided with a back plate, as a main body thereof, formed in an arc shape. A front end of the back plate is movably fixed to the turbine case, where the front end is movable in a direction of an axis of the turbine. The rear end of the back plate is immovably fixed to the turbine case. Such a fixation manner allows thermal displacement of the back plate in the axis direction. An inner surface of the back plate is integrally provided with a touching member for touching with rotating turbine blades, which is formed in a honeycomb shape or the like.
The turbine case is provided with a rear case hook and a C-clip having a C-letter sectional shape supports the rear end of the back plate and the rear case hook in a manner that an outer surface of the rear end of the back plate is touched with the inner surface of the rear case hook. Thereby the rear end of the back plate is fixed to the turbine case and an axial clearance therebetween is sealed so that flow of the hot gas therethrough to the turbine case is suppressed.
The shroud segments and the turbine case are such that the turbine case is prevented from being influenced by the hot gas.
A related art is discloses in Japanese Patent Application Laid-open No. H11-62509.